Turbine engine comprising means for heating the air entering the free turbine

ABSTRACT

The invention relates to a turbine engine ( 10 ), particularly for a helicopter, comprising a gas generator ( 12 ) and a free turbine ( 14 ) rotated by a gas flow (F) generated by the gas generator. 
     The invention is characterized in that the turbine engine also comprises heating means ( 26 ) placed between the gas generator and the free turbine ( 14 ), the said heating means ( 26 ) being capable of increasing the temperature of the gas flow (f) generated by the gas generator ( 12 ) and driving the free turbine.

The present invention relates to the field of gas turbines and in particular that of turbine engines for flying machines such as helicopters.

The present invention relates more particularly to a turbine engine, particularly for a helicopter, comprising a gas generator and a free turbine rotated by the gas flow generated by the gas generator, the turbine engine also comprising a stream connecting an outlet of the gas generator to an inlet of the free turbine, through which the gas flow generated by the gas generator travels.

Conventionally, the gas generator comprises at least one compressor and a high-pressure turbine coupled in rotation by means of a shaft, called the gas generator shaft. The operating principle is as follows: the cool air entering the turbine engine is compressed due to the rotation of the compressor before being sent to a combustion chamber where it is mixed with a fuel. The gases burned due to the combustion are then discharged at great speed.

There then occurs a first expansion in the turbine of the gas generator, during which the latter extracts the energy necessary to drive the compressor.

The turbine of the gas generator does not absorb all the energy of the burned gases and the excess energy corresponds to the gas flow generated by the gas generator.

The latter then supplies energy to the free turbine so that there occurs a second expansion in the free turbine which transforms this energy into mechanical energy in order to drive a receiving member, such as the rotor of the helicopter.

Therefore, during a normal acceleration phase of the turbine engine, particularly in flight, the pilot expresses a need for power which leads to an increase in the flow of fuel injected into the combustion chamber, which has the effect of increasing the energy of the gas flow generated and, consequently, the power of the free turbine.

Naturally, the turbine engine is designed to operate within prescribed limits, particularly at a nominal power, the maintenance of the turbine engine within such limits being carried out by acting notably on the flow of fuel injected into the combustion chamber.

However, in certain circumstances, it is necessary to make the turbine engine function beyond its nominal power for a certain time.

Specifically it is known that the engines of a twin turbine helicopter must be capable of supplying the helicopter with the power necessary to land or continue its flight, at least for a certain time, while one of the two turbine engines has ceased to operate, for example because of a failure.

In other words, in this operating regime, commonly called OEI (One Engine Inoperative), the turbine engine that is still operating must be capable of delivering additional power substantially greater than its nominal power.

Conventionally, during an OEI regime, the pilot substantially increases the flow of fuel injected into the combustion chamber of the gas generator, which has the result of increasing the speed of rotation of the shaft of the gas generator and the temperature in the high-pressure turbine.

The increase in speed and temperature causes an increase in the mechanical stresses in the aerofoils of the high-pressure turbine.

Because of the increase in temperature in the high-pressure turbine, and hence the temperature felt by the blades, the latter sustain, during OEI regimes, mechanical deformations that may cause them to break.

The activation of an OEI regime therefore makes it necessary to more frequently replace the parts making up the turbine engine.

Also and above all, the need for an OEI regime conventionally results in an overdimensioning of the engine which imposes surplus consumption of fuel during the use of the normal regimes.

One object of the present invention is to propose a turbine engine, particularly for a helicopter, that is capable of supplying a power that is substantially greater than its nominal power without imposing an overdimensioning of the engine.

The invention achieves its objective due to the fact that the turbine engine according to the present invention also comprises heating means placed between the gas generator and the free turbine, the said heating means being capable of increasing the temperature of the gas flow generated by the gas generator and driving the free turbine which flows in the said stream, the said heating means comprising a combustion chamber connected to the stream via an upstream passageway so as to be able to be supplied by a fraction of the gas flow generated by the gas generator, the combustion chamber also being connected to the stream via a downstream passageway so as to be able to inject a current of hot gas into the stream.

According to the invention, the fraction of flow which supplies the combustion chamber is strictly smaller than the gas flow generated by the gas generator.

Thanks to the invention, the turbine engine may supply a power that is substantially greater than its nominal power, for example but not necessarily during an OEI regime.

Preferably, the heating means, when they are activated, do not substantially increase the speed of the gases leaving the gas generator.

In the particular case of an OEI regime, the power delivered by the engine will hereinafter be called “OEI power”.

The inventors have in fact noted that the provision of heat energy to the gas flow entering the free turbine, particularly by increasing the temperature of the gas flow leaving the gas generator, preferably but not necessarily without substantially modifying the thermodynamic cycle of the gas generator, results in an increase in the power delivered by the turbine engine.

According to a variant, the turbine engine comprises means for causing a variation in the fraction of flow brought to the combustion chamber.

In other words, the turbine engine according to the invention may supply an OEI power without it being necessary to increase the speed of rotation of the shaft of the gas generator.

Therefore, relative to a conventional turbine engine operating in OEI regime, the high-pressure turbine of the turbine engine according to the invention rotates less quickly and heats up less.

It can therefore be understood that, thanks to the invention, the blades of the high-pressure turbine rotor become damaged less quickly than in a conventional turbine engine.

Also, the present invention makes it possible to reduce the size of the gas generator and therefore the fuel consumption for the nominal regimes.

Naturally, the combustion chamber differs from the combustion chamber of the gas generator, so in the rest of the description, any reference to “combustion chamber”, without further detail, must be understood as corresponding to that of the heating means according to the invention. If necessary, mention will be made of the “combustion chamber of the gas generator”.

It is therefore understood that, when the OEI regime is activated, the combustion chamber is activated and supplies heat energy to the gas flow generated by the gas generator thanks to which its temperature increases.

On advantage of using a combustion chamber is the availability of fuel, already necessary for the operation of the gas generator.

Furthermore, the combustion chamber makes it possible to supply heat energy very quickly to the gas flow generated by the gas generator, which constitutes an assurance of security when the OEI regime is activated.

In addition, thanks to the upstream passageway, a fraction of the flow of the gas generated by the gas generator is tapped off, this fraction of flow then being mixed with fuel then burned in the combustion chamber. The combustion produces a current of hot gas which is evidently hotter than the gas flow generated by the gas generator, this current of hot gas then being reinjected into the stream, preferably by means of said downstream passageway, where it mixes with the gas flow generated by the gas generator, after which the temperature of the gas flow which supplies the free turbine is increased.

This increase in temperature therefore occurs without too much disrupting of the gas flow entering the free turbine stage.

As already indicated, according to the invention, the fraction of the gas flow is substantially smaller than the total gas flow generated by the gas generator. In other words, not all the gas flow leaving the gas generator enters the combustion chamber.

Another advantage of the invention is that it makes it possible to dimension the gas generator independently of the constraints imposed by the emergency regimes specific to multi-engine aviation applications.

An additional advantage is that of supplying an additional transitional torque to the free turbine during the transitional phases in order to limit the drop in speed of rotation of the rotor. This principle applies to all types of turbine engines.

The invention will be better understood and its advantages will better appear on reading the following description of an embodiment indicated as a non-limiting example. The description refers to the appended drawing in which:

FIG. 1 represents a helicopter turbine engine according to the first embodiment of the invention furnished with heating means, comprising a combustion chamber placed between the gas generator and the free turbine.

FIG. 1 is a side view in section of a turbine engine 10 according to a first embodiment of the invention, designed in particular to rotate a rotor of a helicopter (not shown here), the turbine engine 10 comprising a gas generator 12 and a free turbine 14 capable of being rotated by a gas flow F generated by the gas generator 12.

The free turbine 14 is mounted on a shaft 16 which transmits the rotary movement to a receiving member such as a main helicopter rotor.

The turbine engine 10 represented in FIG. 1 is of the type with rear motion offtake. Without departing from the context of the present invention, it would be very possible to consider a free turbine turbine engine of the type with front motion offtake with angled transmission via an external shaft or else a free turbine turbine engine of the type with front motion offtake with angled transmission via a coaxial shaft.

The gas generator comprises a rotary shaft 18, called the gas generator shaft, on which a centrifugal compressor 20 and a high-pressure turbine 22, are mounted, and a combustion chamber 24 placed axially between the compressor 20 and the turbine when considering the gas generator 12 in the axial direction of the rotary shaft 18.

The turbine engine 10 has a casing 25 furnished with an air intake 27 through which the cool air enters the gas generator 12.

After it has been taken into the enclosure of the gas generator 12, the cool air is compressed by the compressor 20 which discharges it towards the entrance of the combustion chamber 24 in which it is mixed with fuel.

The combustion that takes place in the combustion chamber 24 causes the burned gas to be discharged to the high-pressure turbine 22, which has the effect of rotating the shaft 18 of the gas generator 12 and, consequently, the compressor 20.

The rotation speed of the shaft 18 of the gas generator 12 is determined by the flow rate of fuel entering the combustion chamber 24.

Despite the extraction of energy by the high-pressure turbine 22, the gas flow F generated by the gas generator has a significant residual energy.

As can be understood with the aid of FIG. 1, the gas flow F flows through a stream 23, connecting an outlet of the gas generator 12 to an inlet of the free turbine 14, which has the effect of causing an expansion in the free turbine 14 leading to starting the rotation of the rotor of the free turbine and consequently that of the shaft 16.

According to the invention, the turbine engine according to the invention comprises heating means 26 which, according to the first embodiment, include an additional combustion chamber 28, separate from the combustion chamber 24 of the gas generator. As can be seen in FIG. 1, the additional combustion chamber 28 is placed between the gas generator 12 and the free turbine 14 and, more precisely between an outlet of the high-pressure turbine 22 and an inlet of the free turbine 14. A combustion chamber with a torus geometry encircling the stream 23 could be used for example.

Hereinafter, the additional combustion chamber 28 will be simply called “combustion chamber”, while the other will be called “combustion chamber of the gas generator”.

As can be understood with the aid of FIG. 1, the combustion chamber 28 is capable of increasing the temperature of the gas flow generated by the gas generator 12 before it enters the free turbine 14.

To do this, the combustion chamber 28 is connected to the stream 23 by means of an upstream passageway 30 so that a fraction f of the gas flow leaving the gas generator 12 can be bled off and brought into the combustion chamber 28 in order to be mixed with fuel therein. In the meaning of the invention, this fraction f of the gas flow is strictly smaller than the total flow F leaving the gas generator.

The combustion of this mixture in the combustion chamber 28 produces a current of hot gas C, whose temperature is substantially greater than that of the gas flow F generated by the gas generator 12.

As can be seen in FIG. 1, the current of hot gas is reinjected into the stream by means of a downstream passage 32 where it mixes with the portion of the gas flow F that has not been bled off. Since the temperature of the hot gas current C is substantially greater than that of the portion of the flow not bled off, it is understood that the mixture leads to an increase in the temperature of the gas flow flowing in the stream 23 just before it enters the free turbine 14.

As has already been mentioned above, the inventors have discovered that this increase in temperature causes an advantageous increase in the power delivered by the turbine engine while the rotation speed of the shaft 18 of the gas generator 12 is advantageously maintained at its nominal rotation speed.

In this case, arrangement is made for the increase in the temperature entering the free turbine 14 to be between 50 and 200° K, preferably between 80 and 150° K. For this temperature range, it has been noted that the turbine engine could achieve an OEI power similar to that of a conventional turbine engine without increasing the rotation speed of the shaft 18 of the gas generator.

Compared with a turbine engine not fitted with heating means, for example the ARRIEL 2C2 of the applicant, analyses have made it possible to observe that, thanks to the invention, during an OEI regime spanning 30 seconds, for which the OEI power supplied is approximately 20% greater than the nominal power of the turbine engine, the temperature gain of the high-pressure turbine is approximately 80° K while the gain in rotation speed of the shaft of the gas generator is approximately 8%.

In other words, during an OEI regime, the high-pressure turbine of the turbine engine according to the invention rotates more slowly and heats up less than that of a turbine engine not fitted with heating means.

Evidently, the parts making up the free turbine 14 must be dimensioned so as to withstand the increase in temperature.

Another advantage of the invention is that it makes it possible to reduce the weight of the turbine engine.

A further advantage of the present invention, already mentioned, is that it makes it possible to reduce the fuel consumption of the turbine engine 10.

Another advantage is that it makes it possible to offer a turbine engine 10 providing an OEI power greater than that of the equivalent conventional turbine engine.

Without departing from the context of the invention, it is possible to activate the combustion chamber 28 outside the OEI regime, for example in order to temporarily supply additional power to the rotor. 

1. Turbine engine, particularly for a helicopter, comprising a gas generator and a free turbine rotated by a gas flow generated by the gas generator, the turbine engine also comprising a stream connecting an outlet of the gas generator to an inlet of the free turbine, through which the gas flow generated by the gas generator travels, wherein the turbine engine also comprises heating means placed between the gas generator and the free turbine, the said heating means being capable of increasing the temperature of the gas flow driving the free turbine which flows in the said stream, the said heating means comprising a combustion chamber connected to the stream via an upstream passageway so as to be able to be supplied by a fraction of the gas flow generated by the gas generator, the combustion chamber also being connected to the stream via a downstream passageway so as to be able to inject a current of hot gas into the stream.
 2. Turbine engine according to claim 1, wherein the said temperature increase is between 50 and 200° K.
 3. Turbine engine according to claim 1, wherein the combustion chamber encircles the stream. 